This patent application is a regular application of provisional patent application Ser. No. 60/057,594, filed Aug. 29, 1997.
This invention relates to the structure of a liquid-fueled rocket engine, and, more particularly, to the joining of the propellant injector and the combustion chamber.
A typical liquid-fueled rocket engine includes a generally cylindrical combustion chamber, with an injector attached to its injector end and a flared nozzle attached to its nozzle end. A liquid propellant including fuel and an oxidizer flows through injector ports in the injector and into the combustion chamber. The propellant is ignited in the combustion chamber. The hot gas resulting from the combustion expands through the nozzle and drives the rocket engine and the attached rocket structure in the direction opposite to that in which the nozzle is pointed.
The wall of the combustion chamber is exposed to high temperature combustion gas during service. The wall is preferably made of a refractory material such as rhenium coated with iridium on the inwardly facing surface. The injector plate is much cooler, and is typically made of titanium. When the rocket engine is fired, there is a large temperature increase from room temperature and a large temperature gradient between the upper end of the combustion chamber and the adjacent injector, through the region where the two are attached.
It is conventional practice to attach the injector plate to the combustion chamber with a flange-and-bolt system or by electron beam welding. The flange-and-bolt system has the disadvantage that the mechanical seal is not well suited to withstand, without leaking, the high temperatures, large temperature change between room temperature and the service temperature, and large temperature gradients during service. Welding of the extremely dissimilar metals is difficult. The high temperatures on the combustion chamber side of the joint and the high temperature gradient through the joint can lead to a premature failure of the joint and a shortening of the life of the engine. Even a small pinhole in the joint can be disastrous, as it results in a back leak of hot combustion gases from the interior of the combustion chamber.
Recent advances in the design of the rocket engine to allow higher-temperture combustion and the use of more powerful propellants have resulted in a even greater temperatures and temperature gradients. The existing attachment structures may be insufficient for operation in this environment. There is therefore a need for an improved approach to the attachment of the injector to the combustion chamber. The present invention fulfills this need, and further provides related advantages.
The present invention provides a rocket engine with an improved attachment between the injector and the combustion chamber. The rocket engine may be operated reliably at higher temperatures than possible with prior attachment procedures. The fabrication approach uses separately known techniques in a new way.
In accordance with the invention, a rocket engine comprises a combustion chamber comprising an annular wall, an injector, and an attachment between the combustion chamber and the injector. The attachment comprises an annular metallic deposit joined to a first region of the wall of the combustion chamber, an annular transition ring structure, a first weldment between a first welded portion of the transition ring structure and the metallic deposit, and a second weldment between a second welded portion of the transition ring structure and the injector. The first region of the wall of the combustion chamber, to which the metallic deposit is joined, is preferably an outer surface of the chamber wall.
The transition ring structure preferably, but not necessarily, includes an annular step collar, an annular adaptor ring, and a braze joint between at least a portion of the step collar and a portion of the adaptor ring. The step collar protects the end of the combustion chamber inner wall adjacent the injector from damage by the combustion gas and also improves the mixing of the propellants after injection. The step collar is brazed to the adaptor ring, which in turn is welded to the injector and to the metallic deposit.
The metallic deposit may be deposited by any operable technique. Preferably, it is deposited by chemical vapor deposition. The result is a good bond between the metallic deposit and the wall of the combustion chamber. Such a good bond is not easily attained for refractory metals such as the rhenium used in high-performance thrust chamber walls. In this case, the metallic deposit is preferably columbium. (As used herein, a metal identified generically includes both the pure metal and its alloys containing at least about 50 percent by weight of the pure metal. Thus, for example, in the Specification and in the Claims, xe2x80x9ccolumbiumxe2x80x9d includes both pure columbium and its alloys.)
The injector is typically titanium (including both pure titanium and its alloys), and the adaptor ring is typically columbium (including both pure columbium and its alloys). The titanium/columbium weld between the injector and the adaptor ring, and the columbium/columbium weld between the adaptor ring and the metallic deposit, are both readily accomplished by electron beam welding.
The present approach therefore allows the fabrication of a rocket engine from difficult-to-join materials. The joints are sound and gas tight, both at low temperatures and at the elevated temperatures achieved during service. The integrity of the joints is not lost upon the rapid heating of the joint and under imposed high thermal gradients.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the invention. The scope of the invention is not, however, limited to this preferred embodiment.